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Lunar Gemini

Manufacturer's Designation: McDonnell-Douglas. Class: Manned. Type: Lunar Lander. Destination: Moon. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

A direct lunar lander design of 1961, capable of being launched to the moon in a single Saturn V launch through use of a 2-man Gemini re-entry vehicle instead of the 3-man Apollo capsule.

Following the preliminary decision to proceed with the lunar orbit rendezvous technique for the Apollo lunar landing mission, there was one final effort to return to the simpler direct landing approach. Lunar orbit rendezvous would require a three-man Apollo capsule, in order for a crew of two to reach the lunar surface while the third crewmember tended the waiting Apollo in lunar orbit. But if the objective was to land two men on the moon, why use the three-man Apollo capsule? Why not use either the two-man Gemini capsule, or a reduced size two-man Apollo-shaped capsule to land directly on the moon? Such a spacecraft could be propelled toward the moon on a single launch of the Saturn C-5 rocket, just like the LOR version.

McDonnell, builders of the Mercury and Gemini spacecraft, were given the contract to make a study of the alternate approach. The result showed it was indeed feasible, at less cost, risk, complexity, and time, then the LOR 3-man Apollo. The only problem was that the existing work done on the North American three-crew Apollo would have to be scrapped, and either a modified Gemini or a new two-crew Apollo developed in its place. Wiesner, Kennedy's science adviser, was an enthusiastic about the approach. NASA, North American, and the competitors for the lunar module contract were distinctly less interested. Webb, the NASA Administrator, finally got the idea spiked once and for all. Although Gemini lunar landers would be advocated again, as rescue vehicles, or whenever Apollo ran into development trouble, they would never get past the very preliminary paper stage.

Major ground rules for the study were:

  • Launch vehicle was the Saturn C-5, injecting 40,800 kg to a lunar transfer orbit.
  • Mission duration was eight days (two and one-half days flight time to moon, one day on lunar surface plus one day contingency, two and one-half days return time plus one day contingency) plus seven days post landing (one day habitable environment plus six days survivable environment ).
  • Both cryogenic and noncryogenic propulsion systems were to be considered. The final configuration used a LOX/LH2 RL10-powered stage for the major lunar crasher stage, and N2O4/MMH storable propellants for the lunar landing and ascent stages.
  • Atmosphere was 5 psi normal, 3.5 psi emergency, 100% oxygen.
  • First operational flight during the first half of calendar year 1967 was assumed.

Three modifications of the Gemini design were considered for the direct lunar landing:

  • Lunar Gemini I used the as-then-configured, 14 day, earth orbital Gemini command module and service equipment with only those changes considered necessary to effect compatibility with the direct flight lunar landing mission. This included the paraglider for landing at an airstrip at the United States on return. A number of methods for providing the crew with a field of view of the lunar surface during the landing maneuver appear feasible and attractive from the standpoint of minimizing changes to the basic Gemini configuration. The most promising of these were: l) use of the existing Gemini window with an erectable external mirror to provide a downward field of view with the crewman lying in the normal position and 2) use of an auxiliary transparent canopy (or use of the Gemini hatch in open position with cabin depressurized). In the latter method, the crewman was in a rotated (semi-prone) position and viewed the lunar surface directly. A combination of (1) and (2) were selected for Lunar Gemini I with the R.H. crewman provided a mirror and the L.H. crewman provided a transparent canopy through which he may observe the lunar surface while rotated in the present Gemini seat. In conjunction with this position, instruments necessary for control of the lunar landing were provided in an extendable panel located in the service module within the crewman's field of view. Necessary controls were provided adjacent to the seat sides.
  • Lunar Gemini II utilized the then-alternate Gemini 26 m diameter single parachute recovery system in lieu of the paraglider system and associated landing gear. The weight and space savings thus effected permitted the installation of improved navigation and telecommunications capability while increasing the margin for potential weight growth. Earth landings were effected in water and, in the event of emergency recovery over land, the crew utilized the ejection seats to separate from the capsule and terminate the descent with personal parachutes. Crew lunar landing vision provisions were the same as described for Lunar Gemini I.
  • Lunar Gemini III was modified to accept a tower-mounted rocket launch escape system in lieu of ejection seats, thus providing an improved launch abort capability. The paraglider and landing gear were replaced by three 22 m diameter parachutes with normal earth recovery being effected over water. Means were provided for emergency earth recovery over land either through bail-out capability with personal parachutes or, by the use of shock attenuating couches to make land impacts in the command module tolerable. The use of positionable shock attenuated couches in lieu of ejection seats permitted the incorporation of a crew "sit-up" lunar landing capability facilitated by a direct view through a large window in the left-hand hatch.

Weight summary was as follows:

  • Command Module: 2387 to 2551 kg
  • Service Module Equipment: 534 to 553 kg
  • Spacecraft Weight Margin ("to make total come out the same"): 465 to 659 kg
  • Service Module: 10,064 kg
  • Terminal Landing Module: 2735 kg
  • Retrograde Module: 24,393 kg
  • Launch Escape System: net 1,179 kg on Configuration 3
  • Landing Gear Fairing: 634 kg
  • Gross Weight at Launch 41,377 kg to 42,556 kg Less (1% effective)
  • Jettisonable Items: 616 to 1740 kg
  • Effective Launch Weight 40,761 kg to 40,816 kg

The major Gemini Guidance and Navigation components utilized were the inertial system and computer. Additions for Lunar Gemini I included an auto sextant and the Apollo tracking and landing radars (total system weight - 155 kg). Additions for Lunar Gemini II and III included the Apollo sextant/telescope, the Apollo tracking and landing radars, and a roll momentum wheel for use during manual navigation (total system weight - 166 kg).

The power system provided was essentially the same as that used in the 14-day Gemini with some off-loading of fuel cell reactants and the addition of increased sequential control provisions. A detailed electrical load analysis indicated that the mission requirements were 660 watts average for Gemini I and 880 watts average for Lunar Gemini II and III. Sufficient fuel was provided for the full 8 day mission, two days of which were contingency.

Conventional spacecraft structures were employed in all modules, following the proven materials and concepts demonstrated in the Mercury and Gemini designs. Primary structure of each module consisted of a semimonocoque shell with reinforcements around cut-outs and fittings to distribute localized loads. Titanium was used as the basic shell material in all modules except the service module where beryllium sheet was used for the structural radiator shell.

Re-entry heat protection was conservatively designed for a shallow long range re-entry or a 20 g structural limit re-entry, whichever resulted in the greater protection requirements. The ablative material was MAC Thermorad Shield S-3 elastomeric composite. Nominal thermophysical properties were used in the calculations and a 1.15 factor was applied to predicted heating rates. The total ablative material design weight was 232 kg for Lunar Gemini I and II, and 221 kg for Lunar Gemini III.

Crew Size: 2. Length: 12.47 m (40.91 ft). Span: 6.61 m (21.68 ft). Habitable Volume: 2.55 m3. Mass: 18,500 kg (40,700 lb).

  • Gemini Lunar RMOther Designations: Reentry Module. Part of: Gemini Lunar Surface Survival Shelter. Class: Manned. Type: Spacecraft Module.

    Calculated mass based on mission requirements, drawing of spacecraft.

    Crew Size: 2. Length: 3.35 m (10.99 ft). Basic Diameter: 2.32 m (7.61 ft). Maximum Diameter: 2.32 m (7.61 ft). Habitable Volume: 2.55 m3. Mass: 2,386 kg (5,260 lb). Structure Mass: 573 kg (1,263 lb). Heat Shield Mass: 221 kg (487 lb). Reaction Control System: 121 kg (266 lb). Recovery Equipment: 124 kg (273 lb). Navigation Equipment: 118 kg (260 lb). Electrical Equipment: 146 kg (321 lb). Communications Systems: 194 kg (427 lb). Crew Seats and Provisions: 413 kg (910 lb). Crew mass: 144 kg (317 lb). Miscellaneous Contingency: 123 kg (271 lb). Environmental Control System: 150 kg (330 lb).

  • Apollo Direct SMManufacturer's Designation: McDonnell. Class: Manned. Type: Spacecraft Module. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

    The Service Module housed the fuel cells, environmental control, and other major equipment items required for the mission. It also provided the propulsion requirements for lunar launch through transearth midcourse corrections. The same basic engine would used as on the terminal landing module with the throttling feature removed and a radiation cooled skirt added to provide an expansion ratio of 40:1.

    Length: 2.60 m (8.50 ft). Basic Diameter: 2.85 m (9.35 ft). Maximum Diameter: 5.89 m (19.32 ft). Mass: 10,624 kg (23,421 lb). Payload: 520 kg (1,140 lb). Main Engine Propellants: N2O4/MMH. Main Engine Propellants: 8,840 kg (19,480 lb). Electrical System: Fuel Cells.

  • Apollo Direct TLMManufacturer's Designation: McDonnell. Class: Manned. Type: Spacecraft Module. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

    Final letdown, translation hover and landing on the lunar surface from 1800 m above the surface was performed by the terminal landing module. Engine thrust could be throttled down to 1546 kgf. The engine was presumed to be based on a reasonable development of any one of three on-going engine programs for Apollo of the time: Aerojet General's 3600 kgf thrust Transtage engine; Reaction Motors' Model TD-294 4500 kgf thrust LEM research engine; or the 4500 kgf thrust engine under investigation by Rocketdyne for possible use in the LEM. The engine in the terminal landing module was ablatively cooled, throttleable, and had an expansion ratio of 20:1.

    Length: 1.98 m (6.49 ft). Basic Diameter: 5.89 m (19.32 ft). Maximum Diameter: 6.61 m (21.68 ft). Mass: 2,773 kg (6,113 lb). Recovery Equipment: 336 kg (740 lb). Main Engine Thrust: 50.121 kN (11,268 lbf). Main Engine Propellants: N2O4/MMH. Main Engine Propellants: 1,533 kg (3,379 lb).

  • Apollo Direct RMManufacturer's Designation: McDonnell. Class: Manned. Type: Spacecraft Module. Nation: USA. Agency: NASA. Manufacturer: McDonnell.

    The retrograde module supplied the velocity increments required during the translunar portion of the mission up to a staging point approximately 1800 m above the lunar surface. Engine thrust could be throttled down to 1201 kgf.

    Length: 5.00 m (16.40 ft). Basic Diameter: 6.61 m (21.68 ft). Maximum Diameter: 6.61 m (21.68 ft). Mass: 24,393 kg (53,777 lb). Main Engine: RL10A-3. Main Engine Thrust: 117.856 kN (26,495 lbf). Main Engine Propellants: LH2/Lox. Main Engine Propellants: 21,560 kg (47,530 lb).


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